This invention relates to the field of ceramic thermal barrier coatings and to abradable ceramics for use in gas turbine seal applications, and more particularly to a method of producing ceramic thermal barrier coatings and abradable seals comprised of multiple layers in which at least one of the layers is porous.
Gas turbine engines are widely used as sources of motive power, and for other purposes such as electric generation and fluid pumping. Gas turbine manufacturers face a constant customer demand for better performance, enhanced efficiency and improved life. One way to improve performance efficiency and performance is to increase operating temperatures. Increasing operating temperatures usually reduces engine life and is effective only within the limits of materials used in the engine.
Current gas turbine engines are predominantly constructed of metallic materials, with nickel base and cobalt base superalloys being widely used in the higher temperature portions of the engine. Such superalloys are currently used in engines at gas temperatures which are very near, and in some cases above, the melting point of the superalloys. Increases in engine operating temperature are not possible without concurrent steps to protect the superalloys from direct exposure to these high gas temperatures at which the materials would otherwise melt. Such steps include the provision of cooling air (which reduces engine efficiency) and the use of insulating coatings.
Insulating ceramic materials, particularly providing these materials in the form of coatings or thermal barrier coatings, are the primary subject of this invention. Such coatings are most commonly composed of ceramic and are commonly applied by plasma spraying or by electron beam vapor deposition. This invention focuses on coatings applied by electron beam vapor deposition, which is described for example in U.S. Pat. Nos. 4,405,659; 4,676,994 and 5,087,477. Exemplary patents which discuss the current state of the art thermal barrier coatings include U.S. Pat. Nos. 4,321,311; 4,405,660; 5,262,245 and 5,514,482.
The most widely used thermal barrier coating for application to rotating components in turbine engines comprises a bond coat material whose composition is described in U.S. Pat. No. 4,419,416, including a thin layer of aluminum oxide on the bond coat and a columnar grain ceramic coating adhered to the aluminum oxide layer as described in U.S. Pat. No. 4,405,659, developed by the assignee of the present invention. Despite the success of this thermal barrier coating and its widespread acceptance there is a desire for advanced thermal barrier coatings, the principle desired enhancement being improved specific thermal insulation properties, i.e., thermal insulation corrected for density.
If a coating with improved density-corrected insulation properties could be developed, such a coating could either be used at the same thickness as that now used commercially to reduce heat flow, thereby allowing for a reduction in cooling air and enabling a corresponding increase in engine efficiency, or could be used at a reduced thickness to provide the same degree of insulation and heat flow but with reduced coating weight. Such weight reductions are significant, especially on rotating components, since the weight of the thermal barrier coating results in centrifugal forces during engine operation of thousands of pounds on a single turbine blade in a large aircraft engine. Reducing blade centrifugal forces has positive implications in the design requirements of engine components associated with the blade, in particular the supporting disc.
Gas turbine efficiency can also be improved by reducing gas leakage. In particular the clearance between the tips of the rotating blade and the surrounding case structure must be minimized. This is commonly accomplished by providing an abradable seal material on the case. In operation the blade tips cut a channel in the abradable, thus reducing gas leakage. See, e.g., U.S. Pat. Nos. 4,039,296 and 5,536,022, which are expressly incorporated herein by reference.
The present invention comprises a layered ceramic material, preferably applied as a coating. Different layers in the structure have different microstructures, with at least one of the layers being relatively dense and lower defect-containing, and another of the layers being less dense and higher defect-containing. The structure of the less dense layer can be modified by heat treatment to provide porosity. Porosity provides reduced thermal conductivity, and for seal applications the porosity also provides improved abradability. The layers are preferably deposited by electron beam physical vapor deposition. The layers are applied under conditions which produce the previously mentioned differences in density and porosity between alternating layers, by altering the temperature of the substrate and material as deposited.
The relatively dense layers are applied by electron beam vapor deposition under conditions which result in the deposition of what those knowledgeable in the physical vapor deposition art refer to as Zone II structures. The less dense layers (i.e. the layers which will become porous) are also applied by EB-PVD and under conditions which result in microstructures that those skilled in the physical vapor deposition art referred to as Zone I structures. As used herein, the term Zone I means a layer having either an as-deposited Zone I structure or an as-deposited Zone I structure which has been heat treated to enhance porosity.
The resultant structure may be heat treated to enhance porosity through sintering which increases pore size and densifies the ceramic portions which surround the pores.
The invention coating finds particular application in the field of gas turbine components. Such components include turbine airfoils (blades and vanes) and abradable seals which are intended to interact with blade tips or knife edge seals to reduce unwanted gas flow.